JPL PUBLICATION 77-70

An InterstellarPrecursor Mission (NASA-CR-156152) Al fIlERSTELLAR PRECURSOR t178-21173 MISSION (Jet Propulsion Lab.) 111 p HC A06/UF A01 CSCL 22A

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JPL PUBLICATION 77-70

An Interstellar Precursor Mission

L. D. J affe C. Ivie J--. Lewis R. Lipes H. N. Norton J. W. Stearns L. D. Stimpson P. Weissman

October 30, 1977

National Aeronautics and Space Administration

Jet Propulsion Laboratory California Institute of Technology Pasadena, California

Prepared Under Contract No NAS 7-100 National Aeronautics and Space Administration

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PREFACE-

The work described in this report was performed by the Earthand Space

Sciences, Systems, Telecommunications Science and Engineering, Control and

Energy Conversion, Applied Mechanics, and Information Systems Divisions of

the Jet Propulsion Laboratory for NASA Ames Research Center under NASA

OAST Program 790, "Space Systems Studies," Stanley R. Sadin, sponsor.

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ABSTRACT

A mntsion out of the planetary system, with launch about the year 2000,

could provide valuable scientific data as well as test some of the technology

for a later mission to another star. A mission to a star is not expected to

be practical around 2000 because the flight time with the technology then

available is expected to exceed 10,000 yr.

Primary scientific objectives for the precursor mission concern

characteristics of the heliopause, the interstellar medium, stellar distances

(by parallax measurements), low energy cosmic rays, interplanetary gas distri

bution, and mass of the solar system. Secondary objectives include investiga

tion of Pluto. Candidate science instruments are suggested.

The mission should extend to 500-1000 AU from the sun. A heliocentric

hyperbolic escape velocity of 50-100 km/s or more is needed to attain this

distance within a reasonable mission duration. The trajectory should be

toward the incoming interstellar wind. For a year 2000 launch, a Pluto encoun

ter can be included. A second mission targeted parallel to the solar axis

would also be worthwhile.

The mission duration is 20 years, with an extended mission to a total

of 50 years. A system using 1 or 2 stages of nuclear electric propulsion was

selected as a possible baseline. The most promising alternatives are ultralight

solar sails or laser sailing, with the lasers in Earth orbit, for example.

The NEP baseline design allows the option of carrying a Pluto orbiter as a

daughter spacecraft.

Within the limited depth of this study, individual spacecraft systems

for the mission are considered, technology requirements and problem areas

noted, and a number of recommendations made for technology study and advanced

development. The most critical technology needs include attainment of 50-yr

spacecraft lifetime and development of a long-life NEP system.

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RECOMMENDATIONS FOR TECHNOLOGY DEVELOPMENT

FOR EXTRAPLANETARY MISSION

To permit an extraplanetary mission such as that described in this

report,to commence about the year 2000, efforts are recommended on the

following topics. In general, a study should be initiated first, followed

by development effort as indicated by the study.

First priority

Starting work on the following topics is considered of first priority,

in view of their importance to the mission and the time required for the

advance development.

1) Design and fabrication techniques that will provide 50-year space

craft lifetime.

2) Nuclear electric propulsion with operating times of 10 years or more at

full power and able to operate at low power levels for attitude control and

spacecraft power to a total of 50 years.

3) Ultralight solar sails, including their impact upon spacecraft and

mission design.

4) Laser sailing systems, including their impact upon spacecraft and

mission design.

5) Detailing and application of spacecraft quality assurance and reli

ability methods utilizing test times much shorter than the intended lifetime.

Second priority

Other topics that will require advance effort beyond that likely without

special attention include:

6) Spacecraft bearings and moving parts with 50-yr lifetime. 2

Neutral gas mass spectrometer for measuring concentrations of 10

7)

0-10- atom/cm3 , with 50-yr lifetime.

8) Techniques to predict long-time behavior of spacecraft materials from

short-time tests.

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9) Compatibility of science instruments with NEP.

10) Methods of calibrating science instruments for 50-yr lifetime.

11) Optical vs. microwave telecommunications with orbiting DSN.

12) Stellar parallax measurements in deep-space.

FOR STAR MISSION

For a star mission, topics which warrant early study include:

13) Antimatter propulsion.

14) Propulsion alternatives for a star mission.

15) Cryogenic spacecraft.

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TABLE OF CONTENTS

Page

PREFACE ----------------------------------------------------- iii

ABSTRACT ------------------------------------------------------ iv

RECOMMENDATIONS FOR TECHNOLOGY DEVELOPMENT-------------------- v For Extraplanetary Mission ------------------------------------ v

First Priority ---------------------------------------------- v Second Priority v---------------------------------------------V

For Star Mission ---------------------------------------------- vi

INTRODUCTION- -------------------------------------------------- I Background- ---------------------------------------------------- I Study Objective -------------------- --------------------------- I Study Scope ------------------------------------------------- I

STUDY APPROACH ----------------------------------------------- 3 Task I 3--------------------------------------------------------3 Task 2 3--------------------------------------------------------3 Star Mission 4--------------------------------------------------4

SCIENTIFIC OBJECTIVES AND REQUIREMENTS ------------------------ 5 Scientific Objectives 5-----------------------------------------5

Primary Objectives 5------------------------------------------5 Secondary Objectives 5----------------------------------------5

Trajectory Requirements 5---------------------------------------5 Scientific Measurements- --------------------------------------- B Heliopause and Interstellar Medium-------------------------- 8 Stellar and Galactic Distance Scale ------------------------- 9 Cosmic Rays 9-------------------------------------------------9 Solar System as a Whole 0-------------------------------------10 Observations of Distant Objects ----------------------------- 10 Pluto 0-------------------------------------------------------10

Gravity Waves --------------------------------------------- 12

Advantages of Using Two Spacecraft ------------------------- 12

Simulated Stellar Encounter --------------------------------- 11

Measurements Not Planned ------------------------------------ 12

Candidate Science Payload ------------------------------------- 13

TRAJECTORIES ---------------------------------------- 14 Units and Coordinate Systems ---------------------------------- 14 Units ------------------------------------------------------- 14 Coordinate Systems ------------------------------------------ 14

Directions of Interest ---------------------------------------- 14 Extraplanetary ---------------------------------------------- 14 Pluto- ------------------------------------------------------- 15

Solar System Escape Trajectories ------------------------------ 15 Launchable Mass 15-----------------------------------------------15

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Page

TRAJECTORIES (continued) Direct Launch from Earth ------------------------------------- 18 Jupiter Assist ----------------------------------------------- 18

Jupiter Powered Flyby ------------------------------------- 25

Powered Solar Flyby -------------------------------------- 28 Low-Thrust Trajectories ---------------------------------- 28

Solar Sailing -------------------------------------------- 30 Laser Sailing ------------------------------------- 30

Laser Electric Propulsion -------------------------- 31

Fusion ------------------------------------------- ----- 32 Antimatter ------------------------------------------------- 32

Solar Plus Nuclear Electric -------------------------------- 33

Jupiter Gravity Assist ------------------------------------- 18

Launch Opportunities to Jupiter ---------------------------- 25 Venus-Earth Gravity Assist ---------------------- 28

Solar Electric Propulsion ---------------- --. --------- 31

Nuclear Electric Propulsion -------------------------------- 32

Low Thrust Plus Gravity Assist-------------------------- 32

Choice of Propulsion ----------------------------------------- 33

MISSION CONCEPT ---------------------------------------------- 38

MASS DEFINITION AND PROPULSION ------------------------------- 40

INFORMATION MANAGEMENT --------------------------------------- 44

Data Transmission Rate --------------------------------------- 46 Telemetry ---------------------------------------------------- 46

Data Coding-Considerations --------------------- 47 Tracking Loop Considerations ---------------- 47

Options -------------------------------- ---------- 50 More Power -------------------------------------------------- 50

Higher Frequencies --------------------------------------- 53

Selection of Telemetry Option .-------------------------------

Data Generation ---------------------------------------------- 44 Information Management System ---------------- 44 Operations ------------------------- ------------ 45

The Telecommunication Model -------------------------------- 47 Range Equation ---------- ----------------- 47

Baseline Design ---------------------------------------------- 49 Parameters of the System ----------------------------------- 49 Decibel Table and Discussion - ----------------- 50

Larger Antennas and Lower Noise Spectral Density----------- 53

Higher Data Rates ------------------------------------ 55 55

RELATION OF THE MISSION TO SEARCH FOR EXTRATERRESTRIAL INTELLIGENCE ----------------------------------------------- 58

TECHNOLOGY REQUIREMENTS AND PROBLEM AREAS -------------------- 59 Lifetime -------------------------------------------- 59 Propulsion and Power ----------------------------------------- 59

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Page

TECHNOLOGY REQUIREMENTS AND PROBLEM AREAS (continued) Propulsion/Science Interface --------------------------------- 60

Interaction of Thrust with Mass Measurements--------------- 61 Interaction of Thrust Subsystem with Particles and.

Fields Measurements -------------------------------------- 61 Interaction of Power Subsystem with Photon Measurements ---- 61

Telecommunications ------------------------------------------- 62 Microwave vs. Optical Telemetry Systems -------------------- 62 Space Cryogenics ------------------------------------------- 63 Lifetime of Telecommunications Components ------------------ 63 Baseline Enhancement vs. Non-Coherent Communication

System --------------------------------------------------- 64 Information Systems ------------------------------------------ 64 Thermal Control ---------------------------------------------- 64 Components and Materials ------------------------------------ 67 Science Instruments ------------------------------------------ 67 Neutral Gas Mass Spectrometer ------------------------------- 69 Camera Field of View vs, Resolution -------- ------- 69

ACKNOWLEDGEMENT ----------------------------------- ---- 71

REFERENCES ---------------------------------------------- 72

APPENDICES --------------------------------------------------- 75

Appendix A - Study Participants ------------------ 76

Appendix B - Science Contributors ---------------------------- 77

Appendix C - Thoughts for a Star Mission Study---------------- 8 Propulsion ------------------------------------------------- 78 Cryogenic Spacecraft --------------------------------------- 79 Locating Planets Orbiting Another Star --------------------- 81

Appendix D - Solar System Ballistic Escape Trajectories ------ 83

TABLES

1. Position of Pluto, 1990-2030 ----------------------------- 16 2. Summary of Solar System Ballistic Escape Trajectories

Initial Condition: Circular Orbit at 1 AU -------------- 17 3. Capabilities of Shuttle with Interim Upper Stage--------- 19 4. Solar System Escape Using Direct Ballistic Launch

from Earth -------------------------------------------- 20 5. Solar System Escape Using Jupiter Gravity Assist --------- 23 6. Jupiter Gravity Assist Versus Launch Energy -------------- 24 7. Jupiter Powered Flyby ------------------------------- 26 8. Launched Mass for 300 kg Net Payload After Jupiter

Powered Flyby ------------------------------------------ 27 9, Powered Solar Flyby -------------------------------------- 29 10, Performance of Ultralight Solar Sails -------------------- 35

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Page

TABLES (continued)

11. Mass and Performance Estimates for Baseline System ...-.. 43 12. Baseline Telemetry at 1000 AU ----------------- 52 13. Optical Telemetry at 1000 AU -------------------------- 54 14. Telemetry Options ------------------------ 56 15. Proposed Data Rates -------------------------------------- 57 16. Thermal Control Characteristics of Extraplanetary

Missions ----------------------------------------------- 66 17. Technology Rdquirements for Components & Materials 68

FIGURES

1. Some Points of Interest on the Celestial Map ------------ 7 2. Geometry of Jupiter Flyby -------------------- 21 3. Solar System Escape with Ultralight Solar Sails ---------- 34 4. Performance of NEP for Solar Escape plus Pluto ----------- 41 5. Data Rate vs. Ratio of Signal Power to Noise Spectral

Power Density ------------------------------------------ 51

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INTRODUCTION

BACKGROUND

Even before the first earth satellites were launched in 1957, there

was popular interest in the possibility of spacecraft missions to other

stars and their planetary systems. As space exploration has progressed

to the outer planets of the solar system, it becomes appropriate to begin

to consider the scientific promise and engineering difficulties of mission

to the stars and, hopefully, their accompanying planets.

In a conference on "Missions Beyond the Solar System", organized by

L. D. Friedman and held at JPL in August 1976, the idea of a precursor

mission out beyond the planets, but not nearly to another star, was

suggested as a means of bringing out and solving the engineering pro

blems that would be faced in a mission to a star. At the same time, it

was recognized that such a precursor mission, even though aimed primarily

at engineering objectives, should also have significant scientific objec

tives.

Subsequently, in November 1976, this small study was initiated to

examine a precursor mission and identify long lead-time technology develop

ment which should be initiated to permit such a mission. This study was

funded by the Study, Analysis, and Planning Office (Code RX) of the NASA

Office of Aeronautics and Space Technology.

STUDY OBJECTIVE

The objective of the study was to establish probable science goals,

mission concepts and technology requirements for a mission extending from

outer regions of the solar system to interstellar flight. An unmanned

mission was intended.

STUDY SCOPE

The study was intended to address science goals, mission concepts,

and technology requirements for the portion of the mission outward from

the outer portion of the planetary system.

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Because of the limited funding available for this study, it was

originally planned that the portion of the mission between the earth

and the outer portion of the planetary system would not be specifically

addressed; likewise, propulsion concepts and technology would not be

included. Problems encountered at speeds approaching that of light were

excluded for the same reason. In the course of the study, it became

clear that these constraints were not critical, and they were relaxed,

as indicated later in this report.

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STUDY APPROACH

The study effort consisted of two tasks. Task 1 concerned science

goals and mission concepts, Task 2 technology requirements.

TASK I

In Task 1, science goals for the mission were to be examined, and

the scientific measurements to be made. Possible relation of the mission

to the separate effort on Search for Extraterrestrial Intelligence was

also to be considered. Another possibility to be examined was that of

using the data, in reverse time sequence, to examine a star and its sur

roundings (in this case, the solar system) as might be done from an

approaching spacecraft.

Possible trajectories would be evaluated with respect to the inter

action of the direction of the outward asymptote and the speed with the

science goals. A very limited examination might be made of trajectories

within the solar system and accompanying propulsion concepts to assess

the feasibility of the outward velocities considered.

During the study, science goals and objectives were derived by series

of conversations and small meetings with a large number of scientists.

Most of these were from JPL, a few elsewhere. Appendix B gives their

names.

The trajectory information was obtained by examination of pertinent

work done in other studies and a small amount of computation carried out

specifically for this study.

TASK 2

In this task, technology requirements that appear to differ signi

ficantly from those of missions within the solar system were to be identi

fied. These would be compared with the projected state-of-the-art for

the year 2000 ± 15. It was originally planned that requirements associated

with propulsion would be addressed only insofar as they interact with power

or other systems.

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This task was carried out by bringing together study team partici

pants from each of the technical divisions of the Laboratory. (Partici

pants are listed in Appendix A-.) Overal-i concepts were developed and

discussed at study team meetings. Each participant obtained inputs from

other members of his division on projected capabilities and development

needed for individual subsystems. These were iterated at team meetings.

In particular, several iterations were needed between propulsion and tra

jectory calculations.

STAR MISSION

Many of the contributors to this study, both scientific and engineer

ing, felt an actual star mission should be considered. Preliminary exam

ination indicated, however, that the hyperbolic velocity attainable for

solar system escape during the time period of interest (year 2000 ± 15)

was of the order of 102 km/s or 3 x 109 km/year. Since the nearest star 013

is at a distance of 4.3 light years or about 4 x 10 km, the mission dur

ation would exceed 10,000 years. This did not seem worth considering for

two reasons.

First, attaining, and especially establishing, a spacecraft life

time of 10,000 years by the year 2000 is not considered feasible. Secondly,

propulsion capability and hence hyperbolic velocity attainable is expected

to increase with time. Doubling the velocity should take not more than

another 25 years of work, and would reduce the mission duration to only

5000 years. Thus, a spacecraft launched later would be expected to arrive

earlier. Accordingly, launch to a star by 2000 ± 15 does not seem reasonable.

For this reason, a star mission is not considered further in the body

of this report. A few thoughts which arose during this study and pertain

to a star mission are recorded in Appendix C. It is recommended that a

subsequent study address the possibility of a star mission starting in

2025, 2050, or later, and the long lead-time technology developments that

,will be needed to permit this mission.

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SCIENTIFIC OBJECTIVES AND REQUIREMENTS

Preliminary examination of trajectory and propulsion possibilities

indicated that a mission extending to distances of some hundred or per

haps a few thousand AU from the sun with a launch around the year 2000

was reasonable. The following science objectives and requirements are

considered appropriate for such a mission.

SCIENTIFIC OBJECTIVES

Primary Objectives

1) Determination of the characteristics of the heliopause, where the

solar wind presumably terminates against the incoming interstellar

medium.

2) Determination of the characteristics of the interstellar medium.

3) Determination of the stellar and galactic distance scale, through

measurements of the distance to nearby stars.

4) Determination of the characteristics of cosmic rays at energies

excluded by the heliosphere.

5) Determination of characteristics of the solar system as a whole,

such as its interplanetary gas distribution and total mass.

Secondary Objectives

i) Determination of the characteristics of Pluto and its satellites

and rings, if any. If there had been a previous mission to Pluto,

this objective would be modified.

2) Determination of the characteristics of distant galactic and extra

galactic objects.

3) Evaluation of problems of scientific observations of another solar

system from a spacecraft.

TRAJECTORY REQUIREMENTS

The primary science objectives necessitate passing through the helio

pause, preferably in a relatively few years after launch to increase the

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reliability of data return. Most of the scientists interviewed preferred

a mission directed toward the incoming interstellar gaswhere the helio

pause is expected to be closest and most well defined. The "upwind"

direction with respect to neutral interstellar gas is approximately R.A.

250', Decl - 160 (Weller and Meier, 1974; Ajello, 1977). (See Fig. 1.

The sun's motion with respect to interstellar charged particles and mag

netic fields is not known.) Presumably any direction within, say, 400

of this would be satisfactory. A few scientists preferred a mission

parallel to the sun's axis (perpendicular to the ecliptic), believing

that interstellar magnetic field and perhaps particles may leak inward

further along this axis. Some planetary scientists would like the mis

sion to include a flyby or orbiter of Pluto, depending on the extent to which

Pluto might have been explored by an earlier mission. Although a Pluto flyby

is incompatible with a direction perpendicular to the ecliptic, it happens

that in the period of interest (arrival around the year 2005) Pluto will

lie almost exactly in the "upwind" direction mentioned, so an "upwind"

trajectory could include a Pluto encounter.

The great majority of scientists consulted preferred a trajectory

that would take the spacecraft out as fast as possible. This would mini

mize time to reach the heliopause and the interstellar medium. Also, it

would, at any time, provide maximum earth-S/C separation as a base for

optical measurements of stellar parallax. A few scientists would like

to have the S/C go out and then return to the solar system to permit

evaluating and testing methods of obtaining scientific data with a

future S/C encountering another solar system. Such a return would,

roughly, halve the duration of the outward portion of the flight for

any fixed mission duration. Also, since considerable propulsive energy

would be required to "stop and turn around", this approach would con

siderably reduce the outward hyperbolic velocity attainable. These two

effects would greatly reduce the maximum distance that could be reached

for a given mission duration.

As a "strawman mission", it is recommended that a no-return trajec

tory with an asymptote near R.A. 2500, Decl -15o and a flyby of Pluto be

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90 I "11

60 - BARNARD'S 380,000 AU

30-

STAR

APEX OF SOLAR MOTION RELATIVE TO NEARBY STARS

YEAR 2000 30 AU

z o

1990-30 30 AU

0 CELESTIAL EQUATOR

uj

0 *

INCOMING INTERSTELLAR GAS R-30GWELLER AND MEIER (1974)2010AJErLLO (1977)

C

-60 -2

-90, 360

2 34 AU

300 240

* - a C N A RNa.CN UR 270,000 AU

180 120 60 0

RIGHT ASCENSION (0)

Fig. I Some Points of Interest on the Celestial Map. From Sergeyevsky (1971) modified.

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considered, with a hyperbolic excess velocity of 40-90 km/s or more.

Higher velocities should be used if practical. Propulsion should be

designed to avoid interference with scientific measurements and should

be off when mass measurements are to be made.

A number of scientific observations (discussed below) would be

considerably improved if two spacecraft, operating simultaneously, were

used, with asymptotic trajectories at approximately right angles to

each other. Thus, use of a second spacecraft, with an asymptotic tra3ec

tory approximately parallel to the solar axis, is worthwhile scientifically.

SCIENTIFIC MEASUREMENTS

Heliopause and Interstellar Medium

Determination is needed of the characteristics of the solar wind

just inside the heliopause, of the heliopause itself, of the accompanying

shock (if one exists), and of the region between the heliopause and the

shock. The location of the heliopause is not known; estimates now tend

to center at about 100 AU from the sun. (As an indication of the uncer

tainty, estimates a few years ago ran as low as 5 AU.)

Key measurements to be made include magnetic field, plasma proper

ties (density, velocity, temperature, composition, plasma waves) and

electric field. Similar measurements, extending to low energy levels,

are needed in the interstellar medium, together with measurements of the

properties of the neutral gas (density, temperature, composition of atomic

and molecular species, velocity) and of the interstellar dust (particle

concentration, particle mass distribution, composition, velocity). The

radiation temperature should also be measured.

The magnetic, electric, and plasma measurements would require only

conventional instrumentation, but high sensitivity would be needed. Plasma

blobs could be detected by radio scintillation of small sources at a wave

length near 1 m. Radiation temperature could be measured with a radiom

eter at wavelengths of 1 cm to 1 m, using a detector cooled to a few

Kelvins. Both in-situ and remote measurements of gas and dust properties

are desirable. In-situ measurements of dust composition could be made

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by an updated version of an impact-ionization mass spectrometer. In-situ

measurements of ions could be made by a mass spectrometer and by a plasma

analyzer. In-situ measurements of neutral gas composition would probably

require development of a mass spectrometer with greater sensitivity and

signal/noise ratio than present instruments. Remote measurements of gas

composition could be made by absorption spectroscopy, looking back toward

the sun, Of particular interest in the gas measurements are the ratios /1HH/H,Hele;tecnetofN,0adifpsbe+H/H2/H+ , and if possible ofD/i He/H, He3/He4; the contents of C, N,,

Li, Be, B; and the flow velocity. Dust within some size range could be

observed remotely by changes in the continuum intensity.

Stellar and Galactic Distance Scale

Present scales of stellar and galactic distance are probably uncer

tain by 20%. This in turn leads to uncertainties of 40% in the absolute

luminosity (energy production), the quantity which serves as the funda

mental input data for stellar model calculations. Uncertainties in galactic

distances make it difficult to provide good input data for cosmological

models.

The basic problem is that all longer-range scales depend ultimately

on the distances to Cepheid variables in nearby clusters, such as the

Hyades and Pleiades. Distances to these clusters are determined by sta

tistical analysis of relative motions of stars within the clusters, and

the accuracy of this analysis is not good. With a baseline of a few

hundred AU between S/C and earth, triangulation would provide the dis

tance to nearby Cepheids with high accuracy. This will require a camera

with resolution of a fraction of an arc second, implying an objective

diameter of 30 cm to 1 m. Star position angles need not be measured

relative to the sun or earth line, but only with respect to distant

stars in the same image frame. To reduce the communications load, only

the pixel coordinates of a few selected objects need be transmitted to

earth.

Cosmic Rays

Measurements should be made of low energy cosmic rays, which the

solar magnetic field excludes from the heliosphere. Properties to be

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measured include flux, spectrum, composition, and direction. Measurements

should be made at energies below 10 MeV and perhaps down to 10 keV or

lower. Conventional instrumentation should be satisfactory.

Solar System as a Whole

Determinations of the characteristics of the solar system as a

whole include measurements of neutral and ionized gas and of dust. Quan

tities to be measured include spatial distribution and the other proper

ties mentioned above.

Column densities of ionized material can be observed by low fre

quency radio dispersion. Nature, distribution and velocity of neutral

gas components and some ions can be observed spectroscopically by fluores

cence under solar radiation. To provide adequate sensitivity, a large

objective will be needed. Continuum observation should show the dust

distribution.

The total mass of the solar system should be measured. This could

be done through dual frequency radio doppler tracking.

Observations of Distant Objects

Observations of more distant objects should include radio astronomy

observations at frequencies below 1 kHz, below the plasma frequency of the

interplanetary medium. This will require a VLF receiver with a very long

dipole or monopole antenna.

Also, both radio and gamma-ray events should be observed and timed.

Comparison of event times on the S/C and at earth will indicate the direc

tion of the source.

In addition, the galactic hydrogen distribution should be observed

by UV spectrophotometry, outside any local concentration due to the sun.

Pluto

If a Pluto flyby is contemplated, measurements should include optical

observations of the planet to determine its diameter, surface and atmos

phere features, and an optical search for and observations of any satellites

or rings. Atmospheric density, temperature and composition should be measured,

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and nearby charged particles and magnetic fields. Surface temperature and

composition should also be observed. Suitable instruments include a TV

camera, infrared radiometer, ultraviolet/visible spectrometer, particles

and fields instruments, infrared spectrometer.

For atmospheric properties, UV observations during solar occultation

(especially for H and He) and radio observations of earth occultation should

be useful.

The mass of Pluto should be measured: radio tracking should provide

this.

If a Pluto orbiter is included in the mission, measurements should

also include surface composition, variable features, rotation axis, shape,

and gravity field. Additional instruments should include a gamma-ray

spectrometer and an altimeter.

Simulated Stellar Encounter

If return to the solar system is contemplated, as a simulation of a

stellar encounter, observations should be made, during approach, of the

existence of possible stellar companions and planets, and later of satel

lites, asteroids, and comets, and of their characteristics. Observations

of neutral gas, dust, plasma, and energetic emissions associated with the

star should be made, and any emissions from planets and satellites. Choice(s)

should be made of a trajectory through the approaching solar system (recog-_

nizing the time-delays inherent in a real stellar mission), the choice(s)

should be implemented, and flyby measurements made.

The approach measurements could probably be made using instruments

aboard for other purposes. For flyby, it would probably be adequate to

use data recorded on earlier missions rather than carry additional instru

ments.

An alternative considered was simulating a stellar encounter by

"looking backwards while leaving the solar system and later replaying

the data backwards". This was not looked on with favor by the scientists

contacted because the technique would not permit making the operational

decisions that would be key in encountering a "new" solar system: locating

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and flying by planets, for example. "Looking backwards" at the solar system

is desired to give solar system data per se, as mentioned above. Stellar

encounter operations are discussed briefly in Appendix C.

Gravity Waves

A spacecraft at a distance of several hundred AU offers an opportunity

for a sensitive technique for detecting gravity waves. All that is needed

is precision 2-way radio doppler measurements between S/C and earth.

Measurements Not Planned

Observations not contemplated include:

1) Detecting the Oort cloud of comets, if it exists. No method of

detecting a previously unknown comet far out from the sun is recog

nized unless there is an accidental encounter. Finding a previously

seen comet when far out would be very difficult because the nrbits

of long-period comets are irregular and their aphelia are hard to

determine accurately; moreover, a flyby, far from the sun, would

tell little about the comet and nothing about the Oort cloud. The

mass of the entire Oort cloud might be detectable from outside, but

the mission is not expected to extend the estimated 50,000 AU out.

If Lyttleton's comet model is correct, a comet accidentally encoun

tered would be revealed by the dust detector.

2) VLBI using an earth-S/C baseline. This would require very high rates

of data transmission to earth, rates which do not appear reasonable.

Moreover, it is doubtful that sources of the size resolved with

this baseline are intense enough to be detected and that the re

quired coherence would be maintained after passage through inhomo

geneities in the intervening medium. Also, with only 2 widely

separated receivers and a time-varying baseline, there would be

serious ambiguity in the measured direction of each source.

Advantages of Using Two Spacecraft

Use of two spacecraft, with asymptotic trajectories at roughly right

angles to each other, would permit exploring two regions of the heliopause

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(upwind and parallel to the solar axis) and provide significantly greater

understanding of its character, including the phenomena occurring near the

magnetic pole direction of the sun. Observations of transient distant

radio and gamma-ray events from two spacecraft plus the earth would permit

location of the source with respect to two axes, instead of the one axis

determinable with a single S/C plus earth.

CANDIDATE SCIENCE PAYLOAD

1) Vector magnetometer

2) Plasma spectrometer

3) Ultraviolet/visible spectrometers

4) Dust impact detector and analyzer

5) Low energy cosmic ray analyzer

6) Dual-frequency radio tracking (including low frequency with high

frequency uplink)

7) Radio astronomy/plasma wave receiver (including VLF; long antenna)

8) 'Massspectrometer

9) Microwave radiometer

10) Electric field meter

11) Camera (aperture 30 cm to 1 m)

12) Gamma-ray transient detector

If Pluto flyby or orbiter is planned:

13) Infrared radiometer

14) Infrared spectrometer

If Pluto orbiter is planned:

15) Gamma-ray spectrometer

16) Altimeter

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TRAJECTORIES

UNITS AND COORDINATE SYSTEMS

Units

Some useful approximate relations in considering an extraplanetary

mission are:

1 AU = 1.5 x 108km

1 light year = 9.5 x 1012 km = 6.3 x 104 AU

1 parsec = 3.1 x 10 1km = 2.1 x 105 AU = 3.3 light years

1 year = 3.2 x 107

- 61 km/s = 0.21 AU/yr = 3.3 x 10 c

where c = velocity of light

Coordinate Systems

For objects out of the planetary system, the equatorial coordin

ate system using right ascension (a) and declination (6) is often more

convenient than the ecliptic coordinates, celestial longitude (X)

and celestial latitude (8). Conversion relations are:

sin S = cos s sin 6- sin s cos 6 sin a

cos 8 sin X = sin c sin 6 +cos 6 cos & sin a

CoS Cos A = Cos & Cos a

where s = obliquity of ecliptic = 23.50

DIRECTIONS OF INTEREST

Extraplanetary

,Most recent data for the direction of the incoming interstellar

neutral gas are:

Weller & Meier (1974):

Right ascension a = 2520

Declination 6 = -15'

Ajello (1977): ° Right ascension a = 252

Declination 6 = -17*

14

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Thus, these 2 data sources are in excellent agreement.

At a = 2500 the ecliptic is about 200S of the equator, so the wind

comes in at celestial latitude of about 4'. Presumably, it is only

a coincidence that this direction lies close to the ecliptic plane.

The direction of the incoming gas is sometimes referred to as the

"apex of the sun's way", since it is the direction toward which the sun

is moving with respect to the interstellar gas. The term "apex", how

ever, conventionally refers to the direction the sun is moving relative

to nearby stars, rather than relative to interstellar gas. These two

directions differ by about 450 in declination and about 200 in right

ascension. The direction of the solar motion with respect to nearby

stars, and some other directions of possible interest, are shown in

Fig. 1.

Pluto

Table 1 gives the position of Pluto for the years 1990 to 2030.

Note that, by coincidence, during 2000 to 2005 Pluto is within a few

degrees of the direction toward the incoming interstellar gas (see Fig. 1).

At the same time it is near its perihelion distance, only 30-31 AU

from the sun.

SOLAR SYSTEM ESCAPE TRAJECTORIES

As a step in studying trajectories for extraplanetary missions, a

series of listings giving distance and velocity vs. time for parabolic

and hyperbolic solar system escape trajectories has been generated. These

are given in Appendix D and a few pertinent values extracted in Table 2.

Note, for example, that with a hyperbolic heliocentric excess velocity

V = 50 km/s, a distance of 213 AU is reached in 20 years and a distance

of 529 AU in 50 years. With V = 100 km/s, these distances would be

doubled approximately.

LAUNCHABLE MASS

Solar system escape missions typically require high launch energies,

referred to as C3, to achieve either direct escape or high flyby velocity

15

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TABLE 1

Position of Pluto, 1990-2030

Position on 1 January

Distance Right Declination, from ascension, sun,

Year AU 0 0

1990 29.58 227.03 -1.37 1995 29.72 238.51 -6.30 2000 30.12 249.98 -10.89 2005 2010

30.78 31.64

261.39 272.61

-14.92 -18.20

2015 32.67 283.53 -20.69 2020 33.81 294.02 -22.37 2025 35.04 304.00 -23.32 2030 36.31 313.37 -23.63

16

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TABLE 2

Summary of Solar System Ballistic Escape Trajectories

Initial Condition: Circular Orbit at 1 AU

V. Distance (RAD), AU, Velocity (VEL), la/s for Time (T) = for Time (T) =

km/s 10 yrs. 20 yrs. 50 yrs. 10 yrs. 20 yrs. 50 yrs.

0 25.1 40.4 75.3 8.4 6.6 4.9 1 25.2 40.6 76.0 8.4 6.7 4.9 5 27.0 45.1 90.4 9.5 8.0 6.7

10 32.1 57.0 126. 12.5 11.4 10.7 20 47.7 91.2 220. 20.9 20.5 20.2 30 66.5 130. 321. 30.4 30.2 30.1 40 86.5 171. 424. 40.3 40.1 40.1 50 107. 213. 529. 50.2 50.1 50.0 60 128. 254. 634. 60.1 60.1 60.0

(See Appendix D for detail)

17

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at a gravity assist planet. Table 3 gives projected C3 capabilities

in (km/s)2 for the three versions of the Shuttle/Interim Upper Stage

assuming net payloads of 300, 400, and 500 kg. It can be seen that as

launched mass increases the maximum launch energy possible decreases.

Conceivably higher C3' are possible through the use of in-orbit

assembly of larger IUS versions, or development of more powerful upper

stages such as the Tug. The range of C3 values found here will be used

in the study of possible escape trajectories given below.

DIRECT LAUNCH FROM EARTH

Direct launch from the Earth to a ballistic solar system escape

trajectory requires a minimum launch energy of 152.2 (cm/s) 2 . Table 4

gives the maximum solar system V obtainable (in the ecliptic plane) and

maximum ecliptic latitude obtainable (for a parabolic escape trajectory)

for a range of possible C3 .

The relatively low V and inclination values obtainable with direct

launch make it an undesirable choice for launching of extra-solar

probes as compared with those techniques discussed below.

JUPITER ASSIST

Jupiter Gravity Assist

Of all the planets, Jupiter is by far the best to use for gravity

assisted solar system escape trajectories because of its intense gravity

field. The geometry of the Jupiter flyby is shown in Figure 2. Assume

that the planet is in a circular orbit about the Sun with orbital

velocity VJh = 13.06 km/s.

The spacecraft approaches the planet with some relative velocity,

Vin directed at an angle 0 to VJh, and departs along Vout after having,

been bent through an angle a. The total bend angle

2 a = 2 arcsin [1/(I + V. r /p)]in p

where r is the closest approach radius to Jupiter and v= GMJ, the P

gravitational mass of Jupiter. Note that VJh Vin and Vout need not all,

18

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TABLE 3

Capabilities of Shuttle with Interim Upper Stage

Launch energy C3,

(km/s) 2

for indicated

payload (kg)

Launch Vehicle 300 400 500

Shuttle/2-stage IUS 95.5 91.9 88.2

Shuttle/3-stage IUS 137.9 131.0 124.4

Shuttle/4-stage IUS 178.4 161.5 148.2

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TABLE 4

Sblar System Escape Using Direct Ballistic Launch from Earth

Launch

energy, C3,

(km/s)2

152.2

155.

160.

165.

170.

175.

Maximum

hyperbolic

excess velocity,

V , in ecliptic

plane

(km/s)

0.00

3.11

5.15

6.57

7.73

8.72

Maximum ecliptic latitude, ?,Max' for parabolic trajectory

(0)

0.00

2.73

4.53

5.80

6.84

7.74

20

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A

/v

/ escape

//

/

~VA h

Fig. 2 Geometry of Jupiter Flyby

21

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be in the same plane, so the spacecraft can approach Jupiter in the

ecliptic plane and be ejected on a high inclination orbit. The helio

centric velocity of the spacecraft after the flyby, Vsh is given by the,

vector sum of VJh and Vou . If this velocity exceeds approximatelyt

1.414 VJh shown by the &ashed circle in Figure 2, the spacecraft

achieves by hyperbolic orbit and will escape the solar system. The

hyperbolic excess velocity is given by V2sh - 2p/r where V here is GMs,

the gravitation mass for the Sun, and r is the distance from the Sun,

5.2 astronomical units. The maximum solar system escape velocity will

be obtained when the angle between V and Vout is zero. This will

necessarily result in a near-zero inclination for the outgoing orbit.

Around this vector will be a cone of possible outgoing escape trajec

tories. As the angle from the central vector increases the hyperbolic

excess velocity relative to the Sun will decrease. The excess velocity

reaches zero (parabolic escape orbit) when the angle between VJh and

Vsh is equal to arc cos [(3 - V2 in/V 2 Jh)/2 "2. This defines then the

maximum inclination escape orbit that can be obtained for a given V.in at Jupiter. Table 5 gives the dependence of solar system hyperbolic

escape velocity on V. and the angle between V and Vs. The maximumin Jh s angle possible for a given V. is also shown.in

For example, for a V. at Jupiter of 10. km/s the maximum inclinain tion obtainable is 31.41', and the solar system escape speed will be

13.03 km/s for an inclination of 100, 10.45 km/s for an inclination of

20'. Note that for V. 's greater than 20 km/s it is possible to ejectin

along retrograde orbits. This is an undesirable waste of energy however.

It is preferable to wait for Jupiter to move 1800 around its orbit when

one could use a direct outgoing trajectory and achieve a higher escape

speed in the same direction.

To consider in more detail the opportunities possible with Jupiter

gravity assist, trajectories have been found assuming the Earth and

Jupiter in circular, co-planar orbits, for a range of possible launch

energy values. These results are summarized in Table 6. Note that the

orbits with C3 = 180 (km/s)2 have negative semi-major axes indicating

that they are hyperbolic. With the spacecraft masses and launch vehicles

discussed above it is thus possible to get solar syst6m escape velocities

22

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TABLE 5

Solar System Escape Using Jupiter Gravity Assist

Approach velocity relative to Jupiter,Vin (km/s): 6.0 10.0 15.0 20.0 25.0 30.0

Angle between outbound heliocentric velocity of S/C, Vsh, and of Jupiter, Jsh Solar system hyperbolic excess velocity, V., (km/s),

(0) for above approach velocity

0.0 4.70 13.81 21.12 27.42 33.28 38.90 5.0 4.01 13.61 21.00 27.32 33.19 38.82

10.0 13.03 20.63 27.02 32.93 38.58 15.0 12.00 20.01 26.53 32.50 38.19 20.0 10.45 19.13 25.85 31.91 37.65 25.0 8.12 17.99 24.97 31.16 36.97 30.0 3.93 16.57 23.91 30.25 36.15 40.0 12.73 21.25 28.01 34.13 50.0 6.47 17.89 25.28 31.69 60.0 13.75 22.13 28.92 70.0 8.32 18.65 25.94 80.0 14.86 22.83

90.0 10.65 19.70

Maximum angle between outbound heliocentric velocity of S/C, Vsh and of Jupiter, Vjh,(), for above,

approach velocity

9.58 31.41 53.53 76.60 103.57 143.56

Note: * indicates unobtainable combination of V. and angle.

in

23

TABLE 6

Jupiter Gravity Assist versus Launch Energy

Launch Energy,

C3 2 (kn/s)

Transfer orbit semi-major axis,

(AU)

Approach velocity relative to Jupiter, Vin (km/s)

Angle between approach velocity and Jupiter heliocentric velocity,a,

((0)

Maximum Maximum Maximum heliocentric inclination bend hyperbolic to ecliptic angle escape for parabolic relative to velocity, trajectory, Jupiter, a, Vw, for Xmax, for

for flyby flyby at flyby at at 1.1 R 1.1 R 1.1 R

(0) (0)

0.0 90.0

100.0 110.0 120.0 130.0 140.0 150.0 160.0 180.0 200.0

3.23 3.82 4.63 5.82 7.74

11.38 20.98

117.43 -33.64 -9.63 -5.71

6.55 9.08

10.98 12.54 13.88 15.07 16.14 17.12 18.03 19.67 21.13

148.96 127.34 119.53 115.10 112.13 109.95 108.27 106.91 105.78 103.99 102.61

153.85

144.10 137.02 131.35 126.59 122.48 118.86 115.61 112.67 107.52 103.11

6.59

12.22 15.38 17.72 19.61 21.21 22.61 23.87 25.01 27.02 28.77

13.66

27.17 35.81 42.69 48.58 53.82 58.61 63.05 67.22 74.99 82.22

D

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on the order of 25 km/s in the ecliptic plane and inclinations up to

about 670 above the ecliptic plane using simple ballistic flybys of

Jupiter. Thus a large fraction of the celestial sphere is available

to solar system escape trajectories using this method.

Jupiter Powered Flyby

One means of improving the performance of the Jupiter flyby is to

perform a maneuver as the spacecraft passes through periapsis at Jupi

ter. The application of this AV deep in the planet's gravitational

potential well results in a substantial increase in the outgoing Vout

and thus the solar system hyperbolic excess velocity V.. This technique

is particularly useful in raising relatively low Vin values incoming

to high outgoing Vout's. Table 7 gives the outgoing Vou t values at

Jupiter obtainable as a function of V. and AV applied at periapsis.in

A flyby at 1.1 R is assumed. The actual Vou t might be fractionally smaller

because of gravity losses and pointing errors but the table gives a

good idea of the degrees of performance improvement possible.

Carrying the necessary propulsion to perform the AV maneuver would

require an increase in launched payload and thus a decrease in maximum

launch energy and V. possible at Jupiter. Table 8 gives the requiredin launched mass for a net payload of 300 kg after the Jupiter flyby, using

a space storable propulsion system with I of 370 seconds, and the sp maximum C3 possible with a Shuttle/4-stage IUS launch vehicle, as a

function of AV capability at Jupiter. These numbers may be combined

with the two previous tables to find the approximate V. at Jupiter and In

the resulting Vout *

Launch Opportunities to Jupiter

Launch opportunities to Jupiter occur approximately every 13 months.

Precise calculations of such opportunities would be inappropriate at

this stage in a study of extra-solar probe possibilities. Because

Jupiter moves about 330 in ecliptic longitude in a 13 month period, and

because the cone of possible escape trajectories exceeds 300 in half

width for V above about 10 km/s, it should be possible to launch out

to any ecliptic longitude over a 12 year period by properly choosing

the launch date and flyby date at Jupiter. With sufficient V the out

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TABLE 7

Jupiter Powered Flyby

Approach velocity relative toJupaie to Outbound velocity relative to Jupiter, Vo,Jupiter, Vin (kn/s), for indicated AV (km/s) applied

at periapsis of 1.1 R.

.50 1.00 1.50 2.00 2.50

6.0 9.66 12.30 14.48 16.38 18.11 8.0 11.03 13.41 15.44 17.25 18.90

10.0 12.57 14.71 16.59 18.29 19.86 12.0 14.22 16.16 17.00 19.50 20.99 14.0 15.96 17.72 19.33 20.83 22.24

16.0 17.76 19.37 20.86 22.37 23.61 18.0 19.59 21.08 22.47 23.80 25.06 20.0 21.46 22.83 24.14 25.39 26.00

26

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TABLE 8

Launched Mass for 300 kg Net Payload

after Jupiter Powered Flyby

AV at Required launched mass

Jupiter for S/C Is = 370 s

p

(km/s) (kg)

.0 300.

.5 428.

1.0 506.

1.5 602.

2.0 720.

2.5 869.

Maximum launch energy, C3 , attainable with shuttle/4-stage IUS

(km/s) 2

178.4

157.4

147.4

137.0

127.2

114.9

27

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high ecliptic latitudes would be available as described in an earlier

section. Flight times to Jupiter will typically be 2 years or less.

Venus-Earth Gravity Assist

One means of enhancing payload to Jupiter is to launch by way of

a Venus-Earth Gravity Assist (VEGA) trajectory. These trajectories 2

launch at relatively low C3 's, 15 - 30 (km/s) , and incorporate gravity

assist and AV maneuvers at Venus and Earth to send large payloads to

the outer planets. The necessary maneuvers add about 2 years to the

total flight time before reaching Jupiter. The extra payload could

then be used as propulsion system mass to perform the powered flyby

at Jupiter. An alternate approach is that VEGA trajectories allow use

of a smaller launch vehicle to achieve the same mission as a direct

trajectory.

POWERED SOLAR FLYBY

The effect of an impulsive delta-V maneuver when the spacecraft is

close to the Sun has been calculated for an extra-solar spacecraft. The

calculations are done for a burn at the perihelion distance of 0.1 AU,

for orbits whose V value before the burn is 0, 5, and 10 lan/s respective

ly. Results are shown in Table 9. It can be seen that the delta-V

maneuver deep in the Sun's potential well can result in a significant

increase in V after the burn, having its greatest effect when the pre

burn V is small.

The only practical means to get 0.1 from the Sun (other than with a

"super sail", discussed below) is a Jupiter flyby at a V relative to

Jupiter of 12 km/s or greater. The flyby is used to remove angular

momentum from the spacecraft orbit, and "dump" it in towards the Sun.

The same flyby used to add energy to the orbit could achieve V of 17

km/s or more without any delta-V, and upwards of 21 km/s with 2.5 km/s

of delta-V at Jupiter. The choice between the two methods will require

considerably more study in the future.

LOW-THRUST TRAJECTORIES

A large number of propulsion techniques have been proposed that do

not depend upon utilization of chemical energy aboard the spacecraft.

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TABLE 9

Powered Solar Flyby

AV Heliocentric hyperbolic excess velocity, V., (km/s), (km/s) after burn 0.1 AU from Sun and initial V as indicated (km/s)

0 5 10

.1 5.16 7,19 11.25

.3 8,94 10.25 13.42

.5 11.55 12.59 15.29

1.0 16,35 17.10 19.19

1.5 20.05 20.67 22.42

2.0 23.17 23.71 25.26

2.5 25.93 26.41 27.82

29

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Among the more recent reviews pertinent to this mission are those

by Forward (1976), Papailiou et al (1975), and James et al (1976). A

very useful bibliography is that of Mallove et al (1977).

Most of the techniques provide relative low thrust and involve long

periods of propulsion. The following paragraphs consider methods that

seem the more promising for an extraplanetary mission launched around

2000.

Solar Sailing

Solar sails operate by using solar radiation pressure to add or

subtract angular momentum from the spacecraft (Garwin, 1958). The

basic design considered in this study is a helio-gyro of twelve

6200-meter mylar strips, spin-stabilized.

According to Jerome Wright (private communication), the sail is

capable of achieving spacecraft solar system escape velocities of 15-20

km/s. This requires spiralling into a close orbit approximately 0.3 AU

from the sun and then accelerating rapidly outward. The spiral-in

maneuver requires approximately one year and the acceleration outward,

which involves approximately 1-1/2 - 2 revolutions about the sun,

takes about 1-1/2 - 2 years, at which time the sail/spacecraft is

crossing the orbit of Mars, 1.5 AU from the sun, on its way out.

The sail is capable of reaching any inclination and therefore any

point of the celestial sphere. This is accomplished by performing a

"cranking" maneuver when the sail is at 0.3 AU from the sun, before

the spiral outward begins. The cranking maneuver keeps the sail in a

circular orbit at 0.3 AU as the inclination is steadily raised. The

sail can reach 900 inclination in approximately one year's time.

Chauncey Uphoff (private communication) has discussed the possi

bility of a super sail capable of going as close as 0.1 AU from the sun,

and capable of an acceleration outward equal to or greater than the

sun's gravitational attraction. Such a sail might permit escape V's

on the order of 100 km/s, possible up to 300 km/s. However, no such

design exists at present and the possibility of developing such a sail

has not been studied.

Laser Sailing

Rather et al (1976) have recently re-examined the proposal (For

ward, 1962, Marx, 1966,Moeckle, 1972) of using high energy lasers, rather

than sunlight, to illuminate a sail. The lasers could be in orbit

30

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around the earth or moon and powered by solar collectors.

Rather et al found that the technique was not promising for star

missions but could be useful for outer planet missions. Based on their

assumptions , a heliocentric escape velocity of 60 Io/s could be reached

with a laser output power of about 30 kW, 100 km/s with about 1500 kW,

and 200 Im/s with 20 MW. Acceleration is about 0.35 g and thrusting

would continue until the S/C was some millions of kilometers from earth.

Solar Electric Propulsion

Solar electric propulsion uses ion engines, where mercury or

other atoms are ionized and then accelerated across a potential gap

to a very high exhaust velocity. The electricity for generating the

potential comes from a large solar cell array on the spacecraft.

Current designs call for a 100 kilowatt unit which is also proposed

for a future comet rendezvous mission. A possible improvement to the

current design is the use of mirror "concentrators" to focus additional

sunlight on the solar cells at large heliocentric distances.

According to Carl Sauer (private communication) the solar powered

ion drive is capable of escape V.'s on the order of 10-15 km/s in the

ecliptic plane. Going out of the ecliptic is more of a problem because

the solar cell arrays cannot be operated efficiently inside about 0.6 AU

from the sun. Thus the solar electric drive cannot be operated

close iiito the sun for a cranking maneuver as can the solar sail.

Modest inclinations can still be reached through slower cranking or the

initial inclination imparted by the launch vehicle.

Laser Electric Propulsion

An alternative to solar electric propulsion is laser electric:

lasers, perhaps in earth orbit, radiate power to the spacecraft, which is

collected and utilized in ion engines. The primary advantage is that

higher energy flux densities at the spacecraft are possible. This would

permit reducing the receiver area and so, hopefully, the spacecraft

weight. To take advantage of this possibility, receivers that can

operate at considerably higher temperatures than present solar cells will

be needed. A recent study by Forward (1975) suggests that a significant

performance gain, as compared to solar electric, may be feasible.

6 2* Rather et al assumed an allowable flux incident on the sail of 10 W/m laser wavelength 0.5 pm, and laser beam size twice-the diffraction limit. For this calculation, 10 km2 of sail area and 20,000 kg total mass were assumed.

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Nuclear Electric Propulsion

Nuclear electric propulsion (NEP) may use ion engines like solar

electric, or, alternatively, magnetohydrodynamic drive. It obtains

electricity from a generator heated by a nuclear fission reactor.

Thus, NEP is not powertlimited by increasing solar distance.

Previous studies indicate that an operational S/C is possible

by the year 2000 with power levels up to a megawatt (electric) or more

(James et al, 1976).

Preliminary estimates were made based on previous calculations for

a Neptune mission. Those indicated that heliocentric escape velocity

of 50-60 km/s can be obtained.

Fusion

With a fusion energy source, thermal energy could be converted to

provide ion or MHD drive and charged particles produced by the nuclear

reaction can also be accelerated to produce thrust.

A look at one fusion concept gave a V of about 70 ku/s. The

spacecraft weight was 3 x 106 kg. Controlled fusion has still to be

attained.

Bussard (1960) has suggested that interstellar hydrogen could be

collected by a spacecraft and used to fuel a fusion reaction.

Antimatter

Morgan (1975, 1976), James et al, (1976), and Massier (1977a and b)

have recently examined the use of antimatter-matter annihilation to

obtain rocket thrust. A calculation based on Morgan's concepts suggests

that a V over 700 Ion/s could be obtained with a mass comparable to

NEP.

Low Thrust Plus Gravity Assist

A possible mix of techniques discussed would be to use a low

thrust propulsion system to target a spacecraft for a Jupiter gravity

assist to achieve a very high V escape. If for example one accelera

ted a spacecraft to a parabolic orbit as it crossed the orbit of

Jupiter, the V. at Jupiter would be about 17.2 km/s. One could usein

gravity assist then to give a solar system escape V. of 24 ku/s in the

ecliptic plane, or inclinations up to about 63' above the plane.

Powered swingby at Jupiter could further enhance both V and inclination.

32

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A second possibility is to use a solar sail to crank the space

craft into a retrograde (1800 inclination) orbit and then spiral out to

encounter Jupiter at a V. of over 26 km/s. This would result inan escape V 's on the order of 30 km/s and inclinations up to 900, thus

covering the entire celestial sphere. Again, powered swingby would

improve performance but less so, because of the high Vin already present.

This method is somewhat limited by the decreasing bend angle possible

at Jupiter as Vin increases. With still higher approach velocities

the possible performance increment from a Jupiter swingby continues

to decrease.

Solar Plus Nuclear Electric

One might combine solar electric with nuclear electric, using solar

first and then, when the solar distance becomes greater and the solar

distance becomes greater and the solar power falls off, switching to NEP.

Possibly the same thrusters could be used for both. Since operating

lifetime of the nuclear reactor can limit the impulse attainable with NEP,

this combination might provide higher V than either solar or nuclear

electric single-stage systems.

CHOICE OF PROPULSION

Of the various propulsion techniques outlined above, the only

ones that are likely to provide solar system escape velocities above

50 km/s utilize either sails or nuclear energy.

The sail technique could be used with two basic options: solar

sailing, going in to perhaps 0.1 AU from the sun, and laser sailing.

In either case, the requirements on the sail are formidable. Figure 3

shows solar sail performance attainable with various spacecraft light

ness factors (ratios of solar radiation force on the S/C at normal in

cidence to solar gravitational force on the SIC). The sail surface

mass/area ratios required to attain various V values are listed in

Table 10. For a year 2000 launch, it may be possible to attain a sail

surface mass/area of 0.3 g/m2 , if the perihelion distance is constrained

to 0.25 AU or more (W. Carroll, private communication). This ratio

corresponds to an aluminum film about 100 nm thick, which would probably

have to be fabricated in orbit. With such a sail, a V of about 120 km/s

might be obtained.

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300

Jg 200-

X= 0

U,

Ln

Uj LU

jLU

z 100II U 0

01 0 0.1 0.2 0.3

PERIHELION RADIUS, AU

Fig. 3 Solar System Escape with Ultralight Solar Sails. Lightness factor X = (solar radiation force on S/C at normal

incidence)/(solar gravitational force on S/C). From C. Uphoff (private communication).

34

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TABLE 10

PERFORMANCE OF ULTRALIGHT SOLAR SAILS

Initial Heliocentric Lightness Sail Sail Perihelion Excess Factor Load/ Surface Distance Velocity, X Efficiency Mass/Area

Vw aFT/1 a F

g/m2 g/m2

AU km/s

0.25 60 0.8 2.0 0.9

0.25 100 1.8 0.85 0.4

0.25 200 5.5 0.3 0.12

0.1 100 0.6 2.7 1.2

0.1 200 2.2 0.7 0.3

0.1 300 5.0 0.3 0.14

Notes:

X = (solar radiation force on S/C at normal (incidence)/(solar gravitational force on S/C)

aT = (total S/C mass)/(sail area)

p = sail efficiency

= includes sail film, coatings, and seams; excludes structuraloF

and mechanical elements of sail and non-propulsive portions

of S/C. Assumed here: 'IF= 0.5 aT; P = 0.9.

Initial orbit assumed: semi-major axis = 1 x 108 1cm. Sail angle optimized for maximum rate of energy gain.

35

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If the perihelion distance is reduced to 0.1 AU the solar radia

tion force increases but so does the temperature the sail must withstand.

With a reflectivity of 0.9 and an emissivity of 1.0 the sail temperature

would reach 470C (740 K), so high temperature material would have to

be used. Further, according to Carroll (ibid), it may never be possible

to obtain an emissivity of 1.0 with a film mass less than 1 g/m2 ,

because of the emitted wavelength/thickness ratio. For such films an

emissivity of 0.5 is probably attainable; this would increase the tempera

ture to over 6000 C (870 K). Carbon films can be considered, but they

would need a smooth highly reflective surface. It is doubtful a sail

surface mass/area less than 1 g/m2 could be obtained for use at 6000 C. This sail should permit reaching V of 110 km/s: no better than for the

0.25 AU design.

For laser sailing, higher reflectivity, perhaps 0.99, can be

attained because the monochromatic incident radiation permits effective

use of interference layers (Carroll, ibid). Incident energy flux

equivalent to 700 "suns" (at 1 AU) is proposed, however. The high

reflectivity coating reduces the absorbed energy to about the level of

that for a solar sail at 0.1 AU, with problems mentioned above. V.'s up

to 200 km/s might be achieved if the necessary very high power lasers

were available in orbit.

-Considering nuclear energy systems, a single NEP stage using

fission could provide perhaps 60 to 100 km/s V. NEP systems have

already been the subject of considerable study and some advanced develop

ment. Confidence that the stated performance can be obtained is there

fore higher than for any of the competing modes. Using 2 NEP stages or

a solar electric followed by NEP, higher V could be obtained: one

preliminary calculation for 2 NEP stages (requiring 3 shuttle launches

or the year 2000 equivalent) gave V = 150 km/s.

The calculation for a fusion propulsion system indicates 30%

spacecraft velocity improvement over fission, but at the expense of

orders of magnitude heavier vehicle. The cost would probably be pro

hibitive. Moreover, controlled fusion has not yet been attained, and

development of an operational fusion propulsion system for a year 2000

launch is questionable. As to collection of hydrogen enroute to refuel

a fusion reactor, this is further in the future and serious question

exists as to whether it will ever be feasible (Martin, 1972, 1973).

An antimatter propulsion system is even more speculative than a

fusion system and certainly would not be expected by 2000. On the other

36

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hand, the very rough calculations indicate an order of magnitude velocity

improvement over fission NEP without increasing vehicle mass. Also,

the propulsion burn time is reduced by an order of magnitude.

On the basis of these considerations, a fission NEP system was

selected as baseline for the remainder of the study. The very light

weight solar sail approach and the high temperature laser sail approach

may also be practical for a year 2000 mission and deserve further

study. The antimatter concept is the most "far out", but promises orders

of magnitude better performance than NEP. Thus, in future studies

addressed to star missions, antimatter propulsion should certainly be

considered, and a study of antimatter propulsion per se is also warranted.

37

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MISSION CONCEPT

The concept which evolved as outlined above is for a mission out

ward to 500-1000 AU, directed toward the incoming interstellar gas.

Critical science measurements would be made when passing through the

heliopause region and at as great a range as possible thereafter. The

location of the heliopause is unknown but is estimated as 50-100 AU.

Measurements at Pluto are also desired. Launch will be nominally in

the year 2000.

The maximum spacecraft lifetime considered reasonable for a year

2000 launch is 50 years. (This is discussed further, below). To

attain 500-1000 AU in 50 years requires a heliocentric excess velocity

of 50-100 km/s. The propulsion technique selected as baseline is NEP

using a fission reactor. Either 1 or 2 NEP stages may be used. If 2 NEP

stages are chosen, the first takes the form of an NEP booster stage and the

second is the spacecraft itself. The spacecraft, with or without an NEP

booster stage, is placed in low earth orbit by some descendant of the

Shuttle. NEP is then turned on and used for spiral earth escape. Use of

boosters with lower exhaust velocity to go to high earth orbit or earth

escape is not economical. The spiral out from low earth orbit to earth

escape uses only a small fraction of the total NEP burn time and NEP pro

pellant.

After earth escape, thrusting continues in heliocentric orbit. A

long burn time is needed to attain the required velocity: 5 to 10 years are

desirable for single stage NEP (see below), and more than 10 years if two

NEP stages are used. The corresponding burnout distance, depending on the

design, may be as great as 200 AU or even more. Thus, propulsion may be on

past Pluto (31 AU from the sun in 2005) and past the heliopause. To measure

the mass of Pluto, a coasting trajectory is needed; thrust would have to be

shut off temporarily during the Pluto encounter. The reactor would continue

operating at a low level during the encounter to furnish spacecraft power.

Attitude control would preferably be by momentum wheels to avoid any distur

bance to the mass measurements. Scientific measurements, including imagery,

would be made during the fast flyby of Pluto.

38

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After the Pluto encounter, thrusting would resume and continue until

nominal thrust termination ("burnout") of the spacecraft. Enough propel

lant is retained at spacecraft burnout to provide attitude control (unload

ing the momentum wheels) for the 50 year duration of the extended mission.

At burnout the reactor power level is reduced and the reactor provides

power for the spacecraft, including the ion thrusters used for attitude control.

A very useful add-on would be a Pluto Orbiter. This daughter spacecraft

would be separated early in the mission, at approximately the time solar

escape is achieved. Its flight time to Pluto would be about 12 years and its

hyperbolic approach velocity at Pluto about 8 km/s.

The orbiter would be a full-up daughter spacecraft, with enough chemical

propulsion for midcourse, approach, and orbital injection. It would have a

full complement of science instruments (including imaging) and RTG power

sources, and would communicate directly to Earth.

Because the mass of a dry NEP propulsion system is much greater than

that required for the other spacecraft systems, the added mass of a daughter

S/C has relatively little effect on the total inert mass and therefore relatively

little effect on propulsive performance. The mother NEP spacecraft would fly by

Pluto 3 or 4 years after launch, so the flyby data will be obtained at least

5 years before the orbiter reaches Pluto. Accordingly, the flyby data can be

used in selecting the most suitable orbit for the daughter-spacecraft.

If a second spacecraft is to be flown out parallel to the solar axis,

it could be like the one going toward the incoming interstellar gas, but

obviously would not carry an orbiter. Since the desired heliocentric escape

direction is almost perpendicular to the ecliptic, somewhat more propulsive

energy will be required than for the S/C going upwind, if the same escape

velocity is to be obtained. A Jupiter swingby may be helpful. An NEP booster

stage would be especially advantageous for this mission.

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MASS DEFINITION AND PROPULSION

The NEP system considered is similar to those discussed by Pawlik and

Phillips (1977) and by Stearns (1977). As a first rule-of-thumb approximation

the dry NEP system should be approximately 30-35 percent of the spacecraft

mass. A balance is then required between the net spacecraft and propellant,

with mission energy and exhaust velocity being variable. For the very high

energy requirements of the extraplanetary mission, spacecraft propellant

expenditure of the order of 40-60 percent may be appropriate. A booster

stage, if required, may use a lower propellant fraction, perhaps 30 percent.

Power and propulsion system mass at 100-140 km/s exhaust velocity will

be approximately 17 kg/kWe. This is based on a 500 kWe system with 20% con

version efficiency and ion thrusters. Per unit mass may decrease slightly

at higher power levels and higher exhaust velocity. Mercury propellant is

.desired because of its high liquid density, - 13.6 g/cm3 or 13,600 kg/m3

Mercury is also a very effective gamma shield. If an NEP booster is to be

used, it is assumed to utilize two 500 We units.

The initial mass in low earth orbit (M ) is taken as 32,000 kg for

the spacecraft (including propulsion) and as 90,000 kg for the spacecraft

plus NEP booster. 32,000 kg is slightly heavier than the 1977 figure for

the capability of a single shuttle launch. The difference is considered

unimportant, because 1977 figures for launch capability will be only of

historical interest by 2000. 90,000"kg for the booster plus S/C would re

quire the year 2000 equivalent of three 1977 Shuttle launches.

Figure 4 shows the estimated performance capabilities of the propulsion

system for a single NEP stage.

A net spacecraft mass of approximately 1200 kg is assumed and may be broken

out in many ways. Communication with Earth is a part of this and may trade off

with on-board automation, computation and data processing. Support structure for

launch of daughter spacecraft may be needed. Adaptive science capability is also

possible. The science instruments may be of the order of 200-300 kg (including

a large telescope) and utilize 200 kg of radiation shielding (discussed below)

and in excess of 100 W of power. Communications could require as much as 1 kW.

40

140 140

120 120

"- w

0 u

-J0 a W=\

LU<

HV

"

70

60

_u

>

--Ve = 100, Msc = 0

Ve = 120, M = 2000

= 140, Mps/c = 4000 e ' ~V00 FORZ

= =0x 120, Mpc = 2000

e = Vze = 140, Mps/c = 4000

80

-60

,,

U

<

z 0

LU

z

U o -J

40

20-

LUn

40

- 20

3:

0 0 2 4 6

NET SPACECRAFT

8 10 12

MASS (EXCLUDING PROPULSION),

14

kg x 10 3 16

0 18

Fig. 4 Performance of NEP for Solar Escape plus Pluto. 17 kg/kWe.

Ratio (propulsion system dry mass less tankage)/(power input to thrust subsystem) =

a=

= 32,000 kg.

M O = initial mass (in low Earth orbit)

Mps/c = mass of a Pluto S/C separated when heliocentric escape velocity is attained (kg).

Ve = exhaust velocity (km/s).

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One to two kWe of auxiliary power is a first order assumption.

The Pluto Orbiter mass is taken as 500 kg plus 1000 kg of chemical

propellant. This allows a total AV of approximately 3500 m/s and should

permit a good capture orbit at Pluto.

The reactor burnup is taken to be the equivalent of 200,000 hours at

full power. This will require providing reactor control capability beyond

that in existing NEP concepts. This could consist of reactivity poison

rods or other elements to be removed as fission products build up, together

with automated power system management to allow major improvement in adaptive

control for power and propulsion functions. The full power operating time

is, however, constrained to 70,000 h (approximately 8 yr). The remaining

burnup is on reduced power operation for S/C power and attitude control.

At 1/3 power, this could continue to the 50 yr mission duration.

Preliminary mass and performance estimates for the selected system are

given in Table 11. These are for a mission toward the incoming interstellar

wind. The Pluto orbiter, separated early in the mission, makes very little

difference in the overall performance. The NEP power level, propellant

loading, and booster specific impulse were not optimized in these estimates;

optimized performance would be somewhat better.

According to Table 11, the performance increment due to the NEP booster

is not great. Unless an optimized calculation shows a greater increment,

use of the booster is probably not worthwhile.

For a mission parallel to the solar axis, a Jupiter flyby would permit

deflection to the desired 830 angle to the ecliptic with a small loss in VW.

(The approach V. at Jupiter is estimated to be 23 km/s).

in

42

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TABLE 11

Mass and Performance Estimates for Baseline System

(Isp and propellant loading not yet optimized)

Allocation Mass kg

Spacecraft 1200

Pluto orbiter (optional) 1500

NEP (500 kWe) 8500

Propellant: Earth spiral 2100

Heliocentric 18100

Tankage 600

Total for 1-stage (Mo, earth orbit) 32000

Booster 58000

Total for 2-stage (M , earth orbit) 90000

Performance 1 Stage 2 Stages

Booster burnout: Distance 8 AU

Hyperbolic velocity 25 km/s

Time - 4 yr

Spacecraft burnout: Distance (total) 65 155 AU

Hyperbolic velocity 105 150 km/s

Time (total) 8 12 yr

Distance in: 20 yr 370 410 AU

50 yr 1030 1350 AU

43

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INFORMATION MANAGEMENT

DATA GENERATION

In cruise mode, the particles and fields instruments, if reading

continuously, will generate 1 to 2 kb/s of data. Engineering sensors

will provide less. Spectrometers may provide higher raw data rates

but only occasional spectrometric observations would be needed. Star

TV, if run at 10 frames/day (exposures would probably be several hours)

at 108 b/frame would provide about 10 kb/s on the average. A typical

TV frame might include 10 star images whose intensity need be known

only roughly for identification. Fifteen position bits on each axis

and 5 intensity bits would make 350 b/frame or 0.04 b/s of useful data.

Moreover, most of the other scientific quantities mentioned would be

expected to change very slowly, so that their information rate will be

considerably lower than their raw data rate. Occasional transients

may be encountered, and in the region of the heliopause and shock rapid

changes are expected.

During Pluto flyby, data accumulates rapidly. Perhaps 101 bits,

mostly TV, will be generated. These can be played back over a period

of weeks or months. If a Pluto orbiter is flown, it could generate

1010 b/day-or more: an average of over 100 kb/s.

INFORMATION MANAGEMENT SYSTEM

Among the functions of the information handling system will be

storage and processing of the above data. The system compresses the data,

removing the black sky that will constitute almost all of the raw bits

of the star pictures. It will remove the large fraction of bits that

need not be transmitted when a sensor gives a steady or almost-steady

reading. It will vary its processing and the output data stream to

accommodate transients during heliopause encounter and other unpredic

table periods of high information content.

The spacecraft computers system will provide essential support

to the automatic control of the nuclear reactor. It will also support

control, monitoring, and maintenance of the ion thrusters, and of the

attitude control system, as well as antenna pointing and command process

ing.

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According to James et al (1976),the following performance is pro

jected for a S/C information management system for a year 2000 launch:

Processing rate: 109 instruction/s

Data transfer rate: -i09 b/s

Data storage: 'i014 b

Power consumption: 10 - 100 W

Mass: -30 kg

This projection is based on current and foreseen state of the art

and ignores the possibility of major breakthroughs. Obviously, if

reliability requirements can be met, the onboard computer can provide

more capability than is required for the mission.

The processed data stream provided by the information management

system for transmission to earth is estimated to average 20-40 b/s during

cruise. Since continuous transmission is not expected (see below), the

output rate during transmission will be higher.

At heliosphere encounter, the average rate of processed data is

estimated at 1-2 kb/s.

From a Pluto encounter, processed data might be several times 1010

bits. If these are returned over a 6-month period, the average rate

over these months is about 2 kb/s. If the data are returned over a

4-day period, the average rate is about 100 kb/s.

OPERATIONS

For a mission lasting 20-50 years, with relatively little happen

ing most of the time, it is unreasonable to expect continuous DSN

coverage. For the long periods of cruise, perhaps 8 h of coverage per

month, or 1% of the time ,would be reasonable.

When encounter with the heliopause is detected, it might be possible

to increase the coverage for a while; 8 h/day would be more than ample.

Since the time of heliosphere encounter is unpredictable, this possi

bility would depend on the ability of the DSN to readjust its schedule

quickly in near-real time.

For Pluto flyby, presumably continuous coverage could be provided.

For Pluto orbiters, either 8 or 24 h/day of coverage could be provided

for some months.

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DATA TRANSMISSION RATE

On the basis outlined above, the cruise data, at 1% of the time,

would be transmitted at a rate of 2-4 kb/s.

If heliopause data is merely stored and transmitted the same 1% of

the time, the transmission rate rises to 100-200 kb/s. An alternative

would be to provide more DSN coverage once the heliosphere is found.

If 33% coverage can be obtained, the rate falls to 3-7 kb/s.

For Pluto flyby, transmitting continuously over a 6-month period,

the rate is 2 kb/s. At this relatively short range, a higher rate, say,

30-100 kb/s, would probably be more appropriate. This would return the

encounter data in 4 days.

The Pluto orbiter requires a transmission rate of 30-50 kb/s at 24 h/day

or 90-150 kb/s at 8 h/day.

TELEMETRY

The new and unique feature of establishing a reliable telecommunica

tions link for an extraplanetary mission involves dealing with the

enormous distance between the spacecraft (S/C) and the receiving stations

on or near Earth. Current planetary missions involve distances between

the S/C and receiving stations of tens of astronomical units (AU) at

most. Since the extraplanetary mission could extend this distance

to 50